This invention relates generally to gas turbine engine casings and, more particularly, to such structures which are adapted for bleeding interstage air from the compressor. In a gas turbine engine wherein air passes through an inlet to the compressor and hence to a combustion chamber, it is desirable that the thermodynamic conditions of pressure, flow and temperature are uniform about the engine axis through any particular axial position therein. Any distortions of the normal flow pattern through the compressor tends to cause pressure variations across the lateral sections of the engine, thereby resulting in lower efficiency and reduced stall margin. Subsonic aircraft engines in normal flight with normal inlets generally have uniform inlet conditions and, therefore, very little distortion occurs in the airflow pattern. However, in the case of supersonic engines which fly behind supersonic inlets, or subsonic engines which operate within cross wind conditions, distortion of the airflow does tend to occur. This distortion may also occur in aircraft subsonic installations wherein an engine is located in a position such that its axis does not coincide with that of the inlet, as for example in some tail installations where the inlet duct is required to have an "S" shape.
Under the aforesaid conditions, the pressure distortion that occurs is generally highest toward the front of the engine and attenuates as the air moves aft through the engine, but it is not unusual to find substantial pressure variations even as far aft as the combustor.
In order to provide pressurized air for operation of airframe engine accessories such as environmental conditioning, anti-icing, turbine cooling, etc., it is common to include a compressor casing structure which permits bleeding of high pressure air from the compressor to a low pressure plenum. Preferably, this interstage bleeding is accomplished by means which provide minimal interference with the normal airflow patterns in the compressor, but because the manifold provides a communication between areas of high pressure and areas of low pressure, it is possible that air may bleed from one side of the engine to the other side thereof through the manifold. This is particularly true during flight conditions wherein only small amounts of air are being bled from the engine. This communication of air from one side of the engine to the other tends to distort the normal flow pattern in the compressor, or to further the distortion which may be caused by any of the conditions discussed hereinabove.
It is therefore the object of the invention to provide a means of extracting bleed air from an engine that must operate under a variety of pressure distortion conditions in a way that will result in a minimum loss in compressor efficiency and stall margin.
Another object of this invention is to provide in a gas turbine engine a bleed-off system which does not substantially distort the uniform flow of air through the compressor.
Another object of this invention is the provision in a gas turbine engine for an air bleed-off system which operates efficiently over a wide range of flight conditions, wherein varying amounts of air are being bled from the engine.
Another object of this invention is the provision in a gas turbine engine for an air bleed-off manifold which does not allow the air to bleed from one side of the engine to the other through the manifold.
Another object of this invention is the provision for a compressor air bleed-off system which is economical to manufacture and extremely functional in use.
These objects and other features and advantages become more readily apparent upon reference to the following description when taken in conjunction with the appended drawings.